Satellite system

ABSTRACT

A novel cooling system for a superconducting electromagnet (740) that is suitable for use in satellite (700), or at least one or more components of the electromagnet (740) is disclosed. A satellite (700) and electromagnetic control system (705) for position control of such a satellite (700) are also disclosed. In one embodiment, the superconducting magnet control system (705) comprises at least one superconducting electromagnet (740) with at least one cooling element and at least one cryocooler (735). The cryocooler (735) is thermally coupled with the cooling element thereby enabling cooling of the superconducting electromagnet (740) or at least one or more components thereof through the cooling element solely by conduction cooling.

The present invention relates to a satellite system. More particularly but not exclusively it relates to system, method and apparatus for orbiting bodies with a magnetic field.

BACKGROUND TO THE INVENTION

Miniaturized satellites such as CubeSat can be used for space research. CubeSats are made up of multiples of approximately 10 cm×10 cm×10 cm (4 in×4 in×4 in) cubic units. CubeSats have a mass of approximately 1.33 kilograms (2.9 lb) per unit, and often use commercial off-the-shelf (COTS) components for their electronics and structure. CubeSats are commonly put in orbit by deployers on the International Space Station, or launched as payloads on a launch vehicle. Hundreds of CubeSats have been launched already and successfully deployed in the orbit.

Design specifications and basic concepts of such CubeSats can be found in CubeSat website http://www.cubesat.org/, the entirety of which is herein incorporated by reference.

While such miniature satellites are smaller, less costly and have more flexible satellite avionics designs that may be re-usable and easily adaptable across a wide range of satellite missions, several aspects such as cooling and maintaining a thermal mass at cryogenic temperature (below 80 Kelvin), High Temperature Superconducting electromagnet power transfer, integration of components and design of High Temperature Superconducting magnets for use in such satellites have not been sufficiently addressed in the past.

In this specification, where reference has been made to external sources of information, including patent specifications and other documents, this is generally for the purpose of providing a context for discussing the features of the present invention. Unless stated otherwise, reference to such sources of information is not to be construed, in any jurisdiction, as an admission that such sources of information are prior art or form part of the common general knowledge in the art.

For the purpose of this specification, where method steps are described in sequence, the sequence does not necessarily mean that the steps are to be chronologically ordered in that sequence, unless there is no other logical manner of interpreting the sequence.

It is an object of the present invention to provide at least one satellite systems, methods and/or apparatus which overcomes or at least partially ameliorates some of the abovementioned disadvantages or which at least provides the public with a useful choice.

STATEMENTS OF INVENTION

In a first aspect, the present invention may be said to broadly consist in a satellite comprising:

a chassis; and

a superconducting magnet control system mounted to or within at least a portion of the chassis for altitude control or attitude control or both of the satellite, the superconducting magnet control system comprising:

at least one superconducting electromagnet;

at least one power source;

at least one control device;

at least one cryocooling device (preferably) device; and

at least one electromagnetic flux injection device;

the at least one electromagnetic flux injection device being operatively connected to the at least one control device and is configured to derive energy from the at least one power source to energise the superconducting electromagnet(s),

(preferably) the at least one cryocooling device being operatively connected to the at least one control device and configured to derive energy from the at least one power source to cool the at least one superconducting electromagnet.

In one embodiment that satellite further comprises a set of reaction wheels operatively connected to the power and the at least one control device.

In one embodiment the set of reaction wheels comprises three reaction wheels (or three orthogonal reaction wheels).

In one embodiment, the set of reaction wheels derives energy from the at least one power source.

In one embodiment, operation of the set of reaction wheels is controlled by the at least one control device.

In one embodiment, the at least one control device comprises an on-board computer.

In one embodiment, the at least one control device comprises a microprocessor (e.g. a programmable microprocessor).

In one embodiment, the at least one control device comprises a magnetic field sensor (magnetometer).

In one embodiment, the at least one control device comprises a control board.

In one embodiment, the at least one control device derives energy from the at least one power source.

In one embodiment, the at least one control device is configured to control timing and/or magnitude of magnetic field in the electromagnets.

In one embodiment, the at least one control device is configured to control the timing of the cryocooling device to cool the at least one superconducting electromagnet.

In one embodiment, the at least one electromagnetic flux injection device is controlled by the at least one control device to derive energy from the at least one power source to energise the superconducting electromagnet(s).

In one embodiment, the at least one cryocooling device is controlled by the at least one control device to derive energy from the at least one power source to cool the at least one superconducting electromagnet.

In one embodiment, the satellite is configured to be used in a magnetic field (natural or artificial).

In one embodiment, the at least one power source comprises at least one solar panel.

In one embodiment, the at least one power source is a battery.

In one embodiment, the battery is a rechargeable battery.

In one embodiment, the superconducting electromagnet(s) is/are High Temperature Superconductor (HTS) electromagnet(s).

In one embodiment, the electromagnetic flux injection device is an electromagnetic flux pump.

In one embodiment, the electromagnetic flux pump is a linear flux pump.

In one embodiment, the electromagnetic flux pump is contactless.

In one embodiment, the electromagnetic flux pump comprises a plurality of solenoids.

In one embodiment, the electromagnetic flux pump comprises a plurality of solenoids.

In one embodiment, the electromagnetic flux pump comprises a plurality of solenoids that are copper solenoids (or solenoids having copper coil) with an iron core.

In one embodiment, the iron core of each of the plurality of solenoids extend between a first end portion and a second end portion, each first end portion being attached to a common iron frame, wherein, there is a plurality of cubic pieces of iron and each second end portion being attached to one and independent cubic piece from the plurality of cubic pieces.

In one embodiment, the frame is square or is substantially square in shape having a first side, a second side, a third side and a fourth side, the first side being opposite the third side and the second side being opposite the fourth side.

In one embodiment, the electromagnetic flux pump comprises six solenoids, with the first end portion of the iron core of each of three solenoids being attached to a first side of the frame, and the first end portion of the iron core of each of the remaining three solenoids being attached to the third side of the frame, wherein each of the cubic pieces attached to the second end portions are spaced apart from one another.

In one embodiment, the flux pump is a non-linear flux pump.

In one embodiment, the flux pump comprises permanent magnets.

In one embodiment, the flux pump is configured to magnetise HTS coils or HTS coils of tapes.

In one embodiment, the chassis is hollow cubical or hollow cuboid in shape comprising plurality of walls.

In one embodiment, the superconducting electromagnet(s) is/are located in at least one of the walls.

In one embodiment, the superconducting electromagnet(s) is/are High Temperature Superconductor (HTS) electromagnet(s).

In one embodiment, the HTS electromagnet(s) have permeable cores.

In one embodiment, the HTS electromagnet(s) have a relative magnetic permeability greater than that of the conventional cores (e.g. iron cores).

In one embodiment, the HTS electromagnet(s) have a relative magnetic permeability greater than 5000.

In one embodiment, the superconducting electromagnet(s) is/are made out of high temperature superconducting wire double pancake coils.

In one embodiment, the superconducting electromagnet(s) is/are made out of high temperature superconducting wire double pancake coils, based on Yttrium barium copper oxide (YBCO) 2G (second generation).

In one embodiment, the cryocooling device is a cryocooler that is operatively connected (e.g. with a thermal strap) to the superconducting magnet control system (or to the electromagnet(s)) for cooling at least a portion of the superconductive magnet control system.

In one embodiment, the cryocooling device is a tactical integral Stirling or a pulse tube tactical cryocooler that is operatively connected to the superconducting magnet control system for cooling at least a portion of the superconductive magnet control system.

In one embodiment, the cryocooling device is connected to the superconducting magnet with a thermal strap.

In one embodiment, the satellite further comprises a tactical integral Stirling or a pulse tube miniature tactical cryocooler that is operatively connected to the at least one superconducting electromagnet for cooling the superconducting electromagnet(s). In one embodiment, the satellite is a CubeSat comprises at least one unit.

In one embodiment, the satellite dimension of one unit is approximately 100 mm×100 mm×113.5 mm.

In one embodiment, the satellite is approximately 1.33 kg.

In one embodiment, the nominal available power of one unit is approximately between 2 to 10 Watts.

In one embodiment, the maximum magnetic field strength outside a static envelope of the satellite system does not exceed 0.5 Gauss above magnetic field of the Earth.

In one embodiment, the total stored chemical energy of the satellite system does not exceed 100 Watt-Hours.

In one embodiment, the satellite system comprises two or more units that are adjacent to each other.

In one embodiment, the satellite comprises three or more units that are adjacent to each other.

In one embodiment, the chassis comprises a frame structure that is of a four-sided polygonal shape in cross-section, the frame structure being formed by four vertical (or substantially vertical) rails that are spaced apart from each other and four horizontal (or substantially horizontal) rails that are also spaced apart from each other,

wherein, in the frame structure, each of the four vertical rails is connected to two of the other three vertical rails via two of the horizontal rails that are vertically spaced apart from one another.

In one embodiment, the four-sided polygonal shape is a square or a rectangle.

In one embodiment, the vertical rails and the horizontal rails are integrally formed.

In one embodiment, each vertical rail extends between a first end portion and a second end portion, and in each vertical rail one of the horizontal rails is located at or proximal to the first end portion and another one of the horizontal rails is located at or proximal to the second end portion.

In one embodiment, at least a portion of each vertical rail is L-shaped in cross-section.

In one embodiment, the four vertical rails are of same length.

In one embodiment at least one of the vertical rails have plurality of spaced apart apertures along its length.

In one embodiment at least one of the vertical rails have plurality of spaced apart apertures along the length to accommodate countersunk screws.

In one embodiment, the four horizontal rails are of same length.

In one embodiment, the internal volume of the chassis is cubical or cuboid.

In one embodiment, the chassis is constructed of material that is rigid.

In one embodiment, the chassis is constructed of 3D-printed titanium.

In one embodiment, the electromagnet comprises:

a top cooling plate,

a bottom cooling plate, and

a coil sandwiched between the top cooling plate and the bottom cooling plate.

In one embodiment, the coil is a double pancake coil.

In one embodiment, at least one of the top and bottom cooling plates is hexagonal or substantially hexagonal is shape.

In one embodiment, at least one of the top and bottom cooling plates is made from copper.

In one embodiment, at least one of the top and bottom cooling plates is 2 mm in thickness.

In one embodiment, the top and/or bottom cooling plate comprises six holes, preferably, six 3 mm holes on 66 mm diameter.

In one embodiment, the electromagnet comprises a cylindrical magnet bore.

In one embodiment, the cylindrical magnet bore have an outside diameter of about 10 mm and inside diameter of 8 mm.

In one embodiment, the electromagnet comprises two pole pieces with a magnetic field sensor sandwiched between the pole pieces.

In one embodiment, thermal link is provided between top and bottom cooling plates.

In one embodiment, the coil is made from approximately 100 m long, 3 mm wide, 50 μm thick Superpower wire or 2G YBCO HTS tape/wire.

In one embodiment, the coil has 60 mm outer diameter.

In one embodiment, the coil uses approximately 100 m of tape and will be dry-wound with no inter-turn insulation or embedded in a matrix with insulation.

In one embodiment, the inner diameter of the double pancake coils is approximately 10 mm.

In one embodiment, the thickness of the double pancake coils is approximately 3 mm.

In one embodiment, the outer diameter of the double pancake coil is approximately 56 mm.

In one embodiment, the total number of turns of the wire or tape in the double pancake coil is approximately 470.

In one embodiment, the total length of the wire or tape in the double pancake coil is approximately 97.5 m.

In one embodiment, the coil is wrapped around the cylindrical magnet bore.

In one embodiment, insulation sheet (preferably G10 insulation sheet) is provided between the coil (or double pancake coil) and cooling plates to reduce electrical shorting.

In one embodiment, windings of the coil (coil windings) terminates on a current bus.

In one embodiment, the electromagnet comprises yoke plate(s) that is/are attached to exterior surface of one or both of the top or bottom cooling plates.

In one embodiment, thermal grease is to be used/applied in all thermal interfaces.

In one embodiment, the yoke plate is a mild steel yoke plate.

In one embodiment, the yoke plate is a mild steel magnetic yoke plate.

In one embodiment, the yoke plates are screwed or lightly screwed to the top and bottom cooling plates.

In one embodiment, securement means are provided to clamp the top and bottom cooling plates securely for thermal contact with the coil (or double pancake coil).

In one embodiment, securement means are one or more brackets.

In one embodiment, securement means are one or more brackets are made out of stainless steel.

In a second aspect, the present invention may be said to broadly consist a chassis for a satellite comprising a frame structure that is of a four-sided polygonal shape in cross-section, the frame structure being formed by four vertical (or substantially vertical) rails that are spaced apart from each other and four horizontal (or substantially horizontal) rails that are also spaced apart from each other,

wherein, in the frame structure, each of the four vertical rails is connected to two of the other three vertical rails via two of the horizontal rails that are vertically spaced apart from one another.

In one embodiment, the four-sided polygonal shape is a square or a rectangle.

In one embodiment, the vertical rails and the horizontal rails are integrally formed.

In one embodiment, each vertical rail extends between a first end portion and a second end portion, and in each vertical rail, one of the horizontal rails is located at or proximal to the first end portion and another one of the horizontal rails is located at or proximal to the second end portion.

In one embodiment, at least one of the first end portions and the second end portions of each vertical rail comprises a plate member.

In one embodiment, the plate member is integrally formed with the vertical rail.

In one embodiment, at least a portion of each vertical rails is L-shaped in cross-section.

In one embodiment, the four vertical rails are of same length.

In one embodiment at least one of the vertical rails have plurality of spaced apart apertures along the length.

In one embodiment at least one of the vertical rails have plurality of spaced apart apertures along the length to accommodate countersunk screws.

In one embodiment, the four horizontal rails are of same length.

In one embodiment, the internal volume of the chassis is cubical or cuboid.

In one embodiment, the chassis is constructed of material that is rigid.

In one embodiment, the chassis is constructed of 3D-printed titanium.

In one embodiment, the chassis is adapted to be used to support a satellite as defined in any one of the statements of the first aspect.

In a third aspect, the present invention may be said to broadly consist in a satellite comprising a chassis as defined in any one of the statements of the second aspect.

In one embodiment, the satellite is configured to be used in a magnetic field (natural or artificial).

In one embodiment, the satellite is a CubeSat comprising at least one unit.

In one embodiment, the satellite dimension of one unit is approximately 100 mm×100 mm×113.5 mm.

In one embodiment, the satellite is approximately 1.33 kg.

In one embodiment, the nominal available power of one unit is approximately between 2 to 10 Watts.

In one embodiment, the maximum magnetic field strength outside a static envelope of the satellite system does not exceed 0.5 Gauss above magnetic field of the Earth.

In one embodiment, the total stored chemical energy of the satellite system does not exceed 100 Watt-Hours.

In one embodiment, the satellite comprises two or more units.

In one embodiment, the satellite comprises three units.

In a fourth aspect, the invention the present invention may be said to broadly consist in an electromagnet suitable for use in a satellite, the electromagnet comprising:

a top cooling plate,

a bottom cooling plate, and

a coil sandwiched between the top cooling plate and the bottom cooling plate.

In one embodiment, the coil is a double pancake coil.

In one embodiment, at least one of the top and bottom cooling plates are hexagonal or substantially hexagonal is shape.

In one embodiment, at least one of the top and bottom cooling plates is made from copper.

In one embodiment, at least one of the top and bottom cooling plates is 2 mm in thickness.

In one embodiment, the top and/or bottom cooling plate comprises six holes, preferably, six 3 mm holes on 66 mm diameter.

In one embodiment, electromagnet comprises a cylindrical magnet bore.

In one embodiment, the cylindrical magnet bore have an outside diameter of about 10 mm and inside diameter of 8 mm.

In one embodiment, the electromagnet comprises two pole pieces with a magnetic field sensor sandwiched between the pole pieces.

In one embodiment, thermal link is provided between top and bottom cooling plates.

In one embodiment, the coil is made from approximately 100 m long, 3 mm wide, 50 μm thick Superpower wire or 2G YBCO HTS tape/wire.

In one embodiment, the coil has 60 mm outer diameter.

In one embodiment, the coil uses approximately 100 m of tape and is dry-wound with no inter-turn insulation or embedded in a matrix with insulation.

In one embodiment, the inner diameter of the double pancake coils is approximately 10 mm.

In one embodiment, the thickness of the double pancake coils is approximately 3 mm.

In one embodiment, the outer diameter of the double pancake coil is approximately 56 mm.

In one embodiment, the total number of turns of the wire or tape in the double pancake coil is approximately 470.

In one embodiment, the total length of the wire or tape in the double pancake coil is approximately 97.5 m.

In one embodiment, the coil is wrapped around the cylindrical magnet bore.

In one embodiment, the overall size of the electromagnet allows the electromagnet to be able to locate contiguous a wall of a CubeSat.

In one embodiment, insulation sheet (preferably G10 insulation sheet) is provided between the coil (or double pancake coil) and cooling plates to reduce electrical shorting.

In one embodiment, windings of the coil (coil windings) terminates on a current bus.

In one embodiment, the electromagnet comprises a yoke plate(s) that is/are attached to exterior surface of one or both of the top or bottom cooling plates.

In one embodiment, thermal grease is used in all thermal interfaces.

In one embodiment, the thermal grease is Apeizon type N thermal grease.

In one embodiment, the yoke plate(s) is/are a mild steel yoke plate.

In one embodiment, the yoke plate(s) is/are a mild steel magnetic yoke plate.

In one embodiment, the yoke plate(s) is/are screwed or lightly screwed to the top and bottom cooling plates.

In one embodiment, cryocooler is mounted to the electromagnet for cooling.

In one embodiment, the cryocooler is an internal Stirling or a pulse tube miniature tactical cryocooler.

In one embodiment, at least one securement means is provided to clamp the top and bottom cooling plates securely for thermal contact with the coil (or double pancake coil).

In one embodiment, the at least one securement means comprises one or more brackets.

In one embodiment, the at least one securement means is/are one or more brackets made of stainless steel.

In one embodiment, the electromagnet is adapted to be energised by electromagnetic flux injection device.

In one embodiment, the electromagnetic flux injection device is an electromagnetic flux pump.

In one embodiment, the electromagnetic flux pump is a linear flux pump.

In one embodiment, the electromagnetic flux pump is contactless.

In one embodiment, the electromagnet is adapted to be used with the satellite as defined in any one of the above aspects.

In a fifth aspect, the present invention may be said to broadly consist in a satellite comprising an electromagnet as defined in any one statement of the fourth aspect.

In one embodiment, the satellite is configured to be used in a magnetic field (natural or artificial).

In one embodiment, the satellite is a CubeSat comprising at least one unit.

In one embodiment, the satellite dimension of one unit is approximately 100 mm×100 mm×113.5 mm.

In one embodiment, the satellite is approximately 1.33 kg

In one embodiment, the nominal available power of one unit is approximately between 2 to 10 Watts.

In one embodiment, the maximum magnetic field strength outside a static envelope of the satellite system does not exceed 0.5 Gauss above magnetic field of the Earth.

In one embodiment, the total stored chemical energy of the satellite system does not exceed 100 Watt-Hours.

In one embodiment, the satellite comprises two or more units.

In one embodiment, the satellite comprises three units that are adjacent to each other.

In a sixth aspect, the present invention may be said to broadly consist a method of operating a satellite or a satellite system, the method comprising the steps of:

providing a satellite comprising a superconducting magnet control system including at least one superconducting electromagnet, at least one power source, a control device and an electromagnetic flux injection device that are mounted to a portion of a chassis; and

using the control device and the electromagnetic flux injection device to energize the superconducting electromagnet(s) for attitude control or altitude control or both of the satellite or satellite system.

In one embodiment, the method comprises using a cryocooling device to cool the at least one superconducting electromagnet.

In one embodiment, the satellite or satellite system is the one as defined in any one of the above aspects.

In a seventh aspect, the present invention may be said to broadly consist in a method of mounting one or more components in a satellite or satellite system, the method comprising the steps of:

providing a satellite or satellite system having a chassis,

mounting at least one components to a portion of a chassis in such a manner that interior space of the satellite or satellite system is not occupied.

In one embodiment, the satellite or satellite system is the one as defined in any one of the above aspects.

Other aspects of the invention may become apparent from the following description which is given by way of example only and with reference to the accompanying drawings.

In this specification where reference has been made to patent specifications, other external documents, or other sources of information, this is generally for the purpose of providing a context for discussing the features of the invention. Unless specifically stated otherwise, reference to such external documents is not to be construed as an admission that such documents, or such sources of information, in any jurisdiction, are prior art, or form part of the common general knowledge in the art.

For purposes of the description hereinafter, the terms “upper”, “lower”, “right”, “left”, “vertical”, “horizontal”, “top”, “bottom”, “lateral”, “longitudinal” and derivatives thereof shall relate to the invention as it is oriented in the drawing figures. However, it is to be understood that the invention may assume various alternative variations, except where expressly specified to the contrary. It is also to be understood that the specific devices illustrated in the attached drawings, and described in the following description are simply exemplary embodiments of the invention. Hence, specific dimensions and other physical characteristics related to the embodiments disclosed herein are not to be considered as limiting.

It is acknowledged that the term ‘comprise’ may, under varying jurisdictions, be attributed with either an exclusive or an inclusive meaning. For the purpose of this specification, and unless otherwise noted, the term ‘comprise’ shall have an inclusive meaning—i.e. that it will be taken to mean an inclusion of not only the listed components it directly references, but also other non-specified components or elements. This rationale will also be used when the term ‘comprises’ or ‘comprised’ or ‘comprising’ is used in relation to the apparatus or to one or more steps in a method or process.

As used hereinbefore and hereinafter, the term “and/or” means “and” or “or”, or both.

As used hereinbefore and hereinafter, “(s)” following a noun means the plural and/or singular forms of the noun.

When used in claim and unless stated otherwise, the word ‘for’ is to be interpreted to mean only ‘suitable for’, and not for example, specifically ‘adapted’ or ‘configured’ for the purpose that is stated.

This invention may also be said broadly to consist in the parts, elements and features referred to or indicated in the specification of the application, individually or collectively, and any or all combinations of any two or more said parts, elements or features, and where specific integers are mentioned herein which have known equivalents in the art to which this invention relates, such known equivalents are deemed to be incorporated herein as if individually set forth.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will now be described by way of example only and with reference to the drawings in which:

FIG. 1: shows a coordinate system where Points O and P mark the locations of the centres of the magnetic dipoles of the Earth and the on-board magnetic dipole generated by an HTS electromagnet.

FIG. 2: shows a closed loop and an enclosed area used for magnetic moment calculations.

FIG. 3: shows that a Type II superconductor below it's critical temperature expels the applied/external magnetic field from its interior

FIG. 4: shows an example of a 1-unit satellite that is suitable for use in the present invention.

FIG. 5: shows an example of a 2-unit satellite that is suitable for use in the present invention.

FIG. 6: shows an example of a 3-unit satellite that is suitable for use in the present invention.

FIG. 7A: shows a schematic flow diagram of a satellite or satellite system according to one preferred embodiment of the present invention.

FIG. 7B: shows a schematic flow diagram of a satellite or satellite system according to another preferred embodiment of the present invention

FIG. 8: shows panels mounted to a portion of a chassis of the satellite or satellite system for propulsion and/or altitude control. The top and bottom panels and the panels on three sides of the chassis are shown and the panel on one of the side chassis is removed to show the other three panels.

FIG. 9: shows an example of linear flux pump suitable for use in the present invention.

FIG. 10A: shows a front perspective view of a 3-unit chassis according to one preferred embodiment of the present invention for use in satellite or satellite system.

FIG. 10B: shows a top plan view of the chassis of FIG. 9A.

FIG. 10C: shows a rear side elevation view of the chassis of FIG. 9A.

FIG. 10D: show a right-side elevation view of the chassis of FIG. 9A.

FIG. 11A: shows a bottom cooling plate of an electromagnet according to one preferred embodiment of the present invention.

FIG. 11B: shows part of the electromagnet mentioned in FIG. 11A with double pancake coil in place.

FIG. 11C: shows the electromagnet mentioned in FIGS. 11A and 11B with top and bottom cooling plates in place.

FIGS. 11D and E: show the cross section of the electromagnet of FIG. 11C.

FIG. 11F: shows the electromagnet of FIG. 11C with securement means and cryocooler in place.

FIG. 12: shows one example of a cryocooler that may be used in the satellite or satellite system or with electromagnets according to the present invention.

FIGS. 13A-13B: show an experimental set-up of the electromagnet of FIG. 11C.

DETAILED DESCRIPTION

In the description below, unless otherwise stated satellite also refers to a satellite system.

Magnetic propulsion systems are based on the direct interaction of the vehicle's own magnetic field with the natural magnetic field, for instance the geomagnetic one, without using jet propulsion (Pulatov, 2001, the entirety of which is herein incorporated by reference).

The force arising from a dipole-dipole interaction is fundamental to the magnetic propulsion system.

A number of objects in the Solar System (including the Sun) possess a natural magnetic field (Connerney, 1993, Anderson et al., 2011 and T Russell, C Elphic, & Slavin, 1979 the entirety of which is herein incorporated by reference). Some of these fields, subject to their intensity, can be used for satellite propulsion. Low thrust propulsion is achievable in any, however small, external magnetic field. However, strong and dipole-like fields are favourable. The magnetic field of the Earth is understood comparatively well. For this reason, the inventors have utilised its properties to demonstrate how forces can be created and applied to satellite orbit correction.

Although complex, the geomagnetic field can be well approximated by that of a dipole (Fraser-Smith, 1987 the entirety of which is herein incorporated by reference).

For now, it is presumed that the proposed propulsion system, at its simplest form, is constituted by a solenoid, hence giving rise to an on-board magnetic dipole.

The coordinate system is defined in FIG. 1. Points O and P mark the locations of the centres of the magnetic dipoles of the Earth and the on-board solenoid (which is coincident with the on-board dipole) respectively.

Dipolar Coupling

The electrodynamic force arising from the dipole-dipole interaction can be represented by two components: radial FR and tangential Fe. The magnitude of these components can be numerically estimated by Equations 1 and 2 (Pulatov, 2001, the entirety of which is herein incorporated by reference).

$\begin{matrix} {F_{R} = {- \frac{3\mu_{0}M_{E}{M_{P}\left( {{3\mspace{14mu}\cos^{2}\mspace{14mu}\theta} + 1} \right)}^{\frac{1}{2}}}{\left( {4\pi\; R^{4}} \right)}}} & (1) \\ {F_{\theta} = {{- 3}\mu_{0}M_{E}M_{P}\frac{2\mspace{14mu}{\sin\left( {2\theta} \right)}}{\left( {{3\mspace{14mu}\cos^{2}\mspace{14mu}\theta} + 1} \right)^{\frac{1}{2}}\left( {8\pi\; R^{4}} \right)}}} & (2) \end{matrix}$

Where μ_(O) is the permeability of empty space. Distance between the Earth's centre and the centre of the solenoid is denoted by R. Latitude is expressed by θ. Magnetic moments of the Earth and the proposed propulsion system are represented by M_(E) and M_(P), respectively.

The magnetic moment of the propulsion system M_(p) is determined by the properties of the solenoid and its current I, according to Equation 3.

$\begin{matrix} {M_{p} = \frac{ANI}{L}} & (3) \end{matrix}$

Where A is the area enclosed by the solenoid (see FIG. 2) and N is the number of turns.

Magnetic Moment

At all time, the solenoid is subject to an electrodynamic torque, which is described by Equation 4:

τ=M _(P) B ₀ sin α  (4)

Where B_(O) is the magnetic induction vector of the Earth's magnetic field. The angle between vectors B_(O) and M_(p) is denoted by a. It can be observed that the torque is inexistent when the two vectors are aligned (i.e. α=0). This position of perfect alignment shall be referred to as equilibrium further in this specification.

The radial force FR acting on the solenoid in the equilibrium position is directed towards O (the centre of the Earth). However, the direction of the tangential force F_(θ) is position-dependent. It is positive (aligned with the velocity) in quarters I and III (see FIG. 1); negative (opposed to the velocity) in quarters II and IV, or vice-versa, depending on the direction of the orbital motion.

Superconductivity

The phenomenon of superconductivity is utilised to generate stronger on-board magnetic field at a lower energy cost.

It is evident from Equations 1-3 that the magnitude of the generated forces is directly proportional to the magnitude of the electric current I. A high-temperature superconducting wire can facilitate electric current of up to 5.3×104 Ampheres/cm2 (Horvat, Wang, Soltanian, & Dou, 2002, the entirety of which is herein incorporated by reference). Such current is maintained in a solenoid for many years with little or no energy input (Little, 1967, the entirety of which is herein incorporated by reference). To preserve these properties, the wire is kept below its threshold superconducting temperature T_(c), critical current density J_(c) and magnetic field density B.

While superconducting temperature threshold T_(c) is constant and specific to the material choice, other critical values, J_(c) (critical current density) and B_(c) (critical magnetic field density), are dependent on the operative temperature as per Equations 7 and 8 (Dadhich & Schaffner, 2016, the entirety of which is herein incorporated by reference).

$\begin{matrix} {{J_{c}(T)} = {{J_{c}(0)}\left\lbrack {1 - \left( \frac{T}{T_{c}} \right)^{2}} \right\rbrack}} & (7) \\ {{B_{c}(T)} = {{B_{c}(0)}\left\lbrack {1 - \left( \frac{T}{T_{c}} \right)^{2}} \right\rbrack}} & (8) \end{matrix}$

It is evident that the lower the operating temperature T is, the higher current density J_(c) and magnetic field B_(c) can be tolerated by a superconducting wire before its properties vanish.

Additionally, due to Meissner effect, these materials can be used for electromagnetic shielding of the spacecraft's electronics (Hildebrandt, 1970, the entirety of which is herein incorporated by reference). Meissner effect is the expulsion of a magnetic field from the interior of a material that is in the superconducting state, that is, losing its resistance to the flow of electrical currents when cooled below a certain temperature, called the transition temperature. See FIG. 3.

Any further details of how the magnetic propulsion system works will be known to a person skilled in the art and need not be explained in further detail

With the above information, several preferred examples/embodiments of the present invention will now be described.

Reference will now be made to FIG. 4 which shows an example of a 1-unit satellite 400 that is suitable for use in the present invention. The 1-unit satellite 400 comprises six sides namely, top side 441, bottom side 442, left side 446, right side 445, front side 443 and rear side 44. The 1-unit satellite has a standardized form factor bounded by the height h, width w1 and depth w2. By way of example, the height h is about 113.5 mm and the width w1 and the depth w2 are about 100.0 mm. Access port 410 may be provided in the front side 443 and access port 420 may be provided on the rear side 444. The access ports 114 and 116 may provide access to the internal volume of the single unit satellite. There are four rails 430 a-430 d and each end 435 a-d of the rails 430 a-430 d can include one or more deployment switches and/or separation springs that assist in the deployment and/or separation of the satellite 400 from the launch/deployment vehicle. The 1-unit satellite 400 may have a mass of approximately 1.33 kg or more. The nominal available power for one-unit satellite may be between 2 to 10 Watts.

FIG. 5 shows an example of a 2-unit satellite 500 that is suitable for use in the present invention. As it can be seen, satellite 500 comprises two units that are adjacent to each other. This 2-unit satellite 500 has may have the same width w1 and depth w2 as the 1-unit satellite 400 shown in FIG. 4. The 2-unit satellite may have a height h of approximately two times the height h of 1-unit satellite shown in FIG. 1A. Also, there may be two access ports 510 a, 510 b on left hand side and two access ports (only one access port 510 c is shown in FIG. 5) on the right-hand side. The access ports may provide access to the internal volume of the 2-unit satellite 500. Similar to the 1-unit satellite 400, each end 535 a, 535 b, 535 c, 535 d of the rails of this 2-unit satellite can include one or more deployment switches and/or separation springs that assist in the deployment and/or separation of the satellite 500 from the launch/deployment vehicle. Each unit of the satellite 400 may have a mass of approximately 1.33 kg or more. The nominal available power for each unit may be between 2 to 10 Watts.

FIG. 6 shows an example of a 3-unit satellite 600 that is suitable for use in the present invention. As it can be seen, satellite 600 comprises three units that are adjacent to each other. This 3-unit satellite 600 has may have the same width w1 and depth w2 as the 1-unit satellite 400 shown in FIGS. 4 and 2-unit satellite 500 shown in FIG. 5. The 3-unit satellite may have a height h of approximately three times the height of 1-unit satellite shown in FIG. 1A. Also, there may be three access ports 610 a, 610 b and 61 c on left hand side and three access ports (only one access port 610 d is shown in FIG. 6) on the right-hand side. The access ports may provide access to the internal volume of the 3-unit satellite 600. Similar to the 1-unit satellite 400 and 2-unit satellite 200, each end 635 a, 635 b, 635 c, 635 d of the rails of this 3-unit satellite can include one or more deployment switches and/or separation springs that assist in the deployment and/or separation of the satellite 500 from the launch/deployment vehicle. Each unit of the satellite 600 may have a mass of approximately 1.33 kg or more. The nominal available power for each unit may be between 2 to 10 Watts.

In one embodiment, the maximum magnetic field strength outside a static envelope of the satellite 400, 500, 600 may not exceed 0.5 Gauss above magnetic field of the Earth. In one embodiment, total stored chemical energy of the satellite 400, 500, 600 may not exceed 100 Watt-Hours or 100 Watt-Hours per unit.

As mentioned previously, Design specifications and basic concepts of such CubeSats can be found in CubeSat website http://www.cubesat.org/, the entirety of which is herein incorporated by reference.

FIG. 7A shows a schematic flow diagram of a satellite 700 according to one preferred embodiment of the present invention. FIG. 7B shows a schematic flow diagram of a satellite 700 according to another preferred embodiment of the present invention. The satellite 700 shown in FIGS. 7A and 7B are substantially the same except that FIG. 7B also show a set of reaction wheels 745 that the satellite 700 can comprise. There may be three reaction wheels (or three orthogonal reaction wheels) in the set of reaction wheels 745. A skilled person will appreciate that the reaction wheels may be beneficial to the system and may be required for certain functionality. The set of reaction wheels 745 may derive energy from the power source 715. The operation of the at least three reaction wheels may be controlled by the control device 725.

FIG. 8 shows panels mounted to a portion of a chassis. The panels may comprise electromagnet(s) (not shown) for attitude and/or altitude control of the satellite. The top and bottom panels and the panels on three sides of the chassis are shown and the panel on one of the side chassis is removed to show the other three panels. Three orthogonal electromagnets may be required to achieve the greatest control over satellite's position. The electromagnets can be split in 2 (hence resulting in 6 magnets in total) to achieve uniform mass distribution.

The satellite 700 may comprise the chassis 710. A superconducting magnet control system 705 may be mounted to or within at least a portion of the chassis 710 for attitude and/or altitude control of the satellite 700. The superconducting magnet system 705 may comprise at least one power source 715, at least one control device 725, at least one electromagnetic flux injection device 730, at least one control device and at least one superconducting electromagnet 740. The electromagnetic flux injection device 730 may be operatively connected to the control device 725 and may be configured to derive energy from the at least one power source 715 to energise the superconducting electromagnet(s) 740 which then allows for magnetic/electromagnetic attitude control and/or altitude control of the satellite 700.

As shown, the superconducting magnet control system 705 may comprise at least one cryocooling device 725 that may be operatively connected to the control device 725. The cryocooling device 725 may be configured to derive energy from the power source 715 to cool the at least one superconducting electromagnet 740.

As shown in FIG. 7B, the satellite 700 can comprise a set of reaction wheels 745. There may be three reaction wheels (or three orthogonal reaction wheels) in the set of reaction wheels 745. A skilled person will appreciate that the reaction wheels may be beneficial to the system and may be required for certain functionality.

The control device 725 may comprise or be in the form of an on-board computer or a microprocessor (e.g. a programmable microprocessor or a control board). The control device may comprise or be in the form of a magnetic field sensor (magnetometer).

The control device 725 may derive energy from the power source 715.

The control device 725 may be configured to control timing and/or magnitude of magnetic field in the superconducting electromagnet 740.

The control device may also be configured to control the timing of the cryocooling device 735 to cool the at least one superconducting electromagnet 740.

The electromagnetic flux injection device 730 may be controlled by the control device 725 to derive energy from the at least one power source to energise the superconducting electromagnet(s) 740.

Similarly, the cryocooling device 735 may be controlled by the at least one control device 725 to derive energy from the at least one power source 715 to cool the superconducting electromagnet 740.

The satellite 700 may be used in a magnetic field that is artificial and/or natural. The power source 715 may comprises at least one solar panel and/or on-board power supply e.g. a battery or a rechargeable battery. Alternatively, many other form of a suitable power source(es) may be used.

The superconducting electromagnet(s) 735 may be High Temperature Superconductor (HTS) electromagnets 1100 which will be described later with reference to FIGS. 11 A to 11F. The HTS electromagnet(s) 1100 may have a relative magnetic permeability greater than that of the conventional cores (e.g. iron cores). The HTS electromagnet(s) 1100 may have a relative magnetic permeability greater than 5000.

The concept of how the energised superconducting electromagnet(s) 735 allows for the altitude control and/or attitude control of the satellite 700 will be apparent to a person skilled in the art and need not be described in any further detail. A skilled person will appreciate that attitude control can be done with regular electromagnets. This is known in the industry as magnetorquer.

Also, a skilled person will appreciate that the electromagnetic altitude control system may provide 6 degrees of freedom which can allow electromagnetic attitude control that requires three degrees of freedom.

As mentioned above, the present invention may use electromagnetic flux injection device 725 that may be configured to derive energy from at least one power source 715 to energise the superconducting electromagnet(s) 740 such as HTS electromagnet(s). The at least one power source 715 may be a solar panel and/or a battery e.g. a rechargeable battery. As mentioned above, many other forms of a suitable power source(es) may be used. Preferably, the power source 715 comprises both the solar panel and the on-board power supply.

The electromagnetic flux injection device 730 may be an electromagnetic flux pump such as a linear flux pump. HTS coils are not entirely resistance free. Ohmic losses occur at the welding points. Therefore, continuous energy input may be required to maintain a particular magnetic field strength. Such energy input cannot be feasibly implemented with electric current leads (wires attached to a superconducting coil, which supply power) as they can cause further loses. At least for that reason, it is preferred that the electromagnetic injection device 730 is in the form of electromagnetic flux pump that is contactless.

A skilled person will appreciate how an electromagnetic flux injection device 730 or electromagnetic flux pump may be configured to derive energy from at least one power source to energise the superconducting electromagnet(s) such as HTS electromagnet(s), and therefore that need not be described in detail.

A paper concerning a linear flux pump design which could be used to magnetize HTS tapes and coils is described in Fu, L., Matsuda, K., Baghdadi, M., & Coombs, T. (2015). Linear Flux Pump Device Applied to High Temperature Superconducting (HTS) Magnets. IEEE Transactions on Applied Superconductivity, 25(3), 1-4, which is incorporated by reference herein in its entirety. The design is based on an iron magnetic circuit together with copper solenoids and is powered by a current source driver circuit.

FIG. 8 shows an example of linear flux pump 800 (similar to the one as described in that paper) that can be used to energize the superconducting electromagnet(s) 740 such as HTS electromagnet(s).

As it can be seen in FIG. 9, the flux pump 800 may comprise a plurality of solenoids 801-806 that may be copper solenoids (or solenoids having copper coil) with an iron core 811-816. Of course, non-copper solenoids and non-iron cores may be used. Each of the iron cores 811-816 of the solenoids may extend between a first end portion and a second end portion. In this example, there are six solenoids 801-806 each with an iron core 811-816 that extends between a first end portion and a second end portion. The first end portion of each iron core may be attached to a common iron frame 820.

As shown in FIG. 9, the frame 820 may be square or substantially square in shape having a first side 821, a second side 822, a third side 823 and a fourth side 824. In this embodiment of the flux pump 800 comprising six solenoids 801-806, the first end portion of the iron core of each of three solenoids 802, 804, 806 may be attached to a first side 821 of the frame 820, and the first end portion of the iron core of each of the remaining three solenoids 801, 803, 805 may be attached to the third side 823 of the frame. The cubic pieces 831-836 are attached to the second end portions and the cubic pieces 831-836 are spaced apart from one another. As shown, the second portion of each of the iron cores 811-816 is attached to one and independent cubic piece. In other words, as it can be seen, second portion of iron core 811 may be attached to cubic piece 831, second portion of iron core 812 may be attached to cubic piece 832, second portion of iron core 813 may be attached to cubic piece 833, second portion of iron core 814 may be attached to cubic piece 834, second portion of iron core 815 may be attached to cubic piece 835 and second portion of iron core 816 may be attached to cubic piece 836.

The flux pump 800 of FIG. 9 may be configured to magnetise coils of superconducting electromagnet. This should be apparent to a person skilled in the art especially upon considering Fu, L., Matsuda, K., Baghdadi, M., & Coombs, T. (2015). Linear Flux Pump Device Applied to High Temperature Superconducting (HTS) Magnets. IEEE Transactions on Applied Superconductivity, 25(3), 1-4, which is incorporated by reference herein in its entirety.

Instead of the linear flux pump 800, a non-linear flux pump (e.g. non-linear flux pumps using permanent magnets instead of solenoids) may be used.

FIGS. 10A-10D show a lightweight chassis 900 of a satellite, according to one preferred embodiment of the present invention. As shown, chassis 900 has minimal structural support allowing it to offer the space, which would otherwise be dedicated to structural satellite walls. As it can be seen, this design eliminates the chassis walls altogether, and may rely on extremely rigid and lightweight 3D printed titanium. Many other suitable alloys many be used.

The chassis for a satellite may be in a form of a frame structure that is of a four-sided polygonal shape (such as rectangular) in cross-section. The frame structure may be formed by four vertical (or substantially vertical) rails 905 a-d that are spaced apart from each other and four horizontal (or substantially horizontal) rails 910 a-d that are spaced apart from each other. Each one of the four vertical rails 905 a-d may be connected to two other vertical rails via two horizontal rails that vertically spaced apart from one another. For example, as shown in FIG. 10A, vertical rail 905 a is connected to two other vertical rails 905 b and 905 c via two horizontal rails 910 a and 910 d respectively that are vertically spaced apart. Other vertical rails 905C-905D are also each connected to two other vertical rails via two vertically spaced apart horizontal rails in a similar manner as shown.

The vertical rails 905 a-d and the horizontal rails 910 a-d may be integrally formed.

As shown, each vertical rail 905 a-d may extend between a first end portion and a second end portion and in each vertical rail 905 a-d one of the horizontal rails is located at or proximal to the first end portion and another one of the horizontal rails is located at or proximal to the second end portion. For example, in vertical rail 905A, horizontal rails 910 a and 910 d are positioned such that one of the horizontal rails is proximal to the first end portion and the another one is proximal to the second end portion.

At least one of the first end portions and the second end portions of each vertical rail may comprise a cap or a plate member 920. The plate member 920 of may be integrally formed with the vertical rail(s) 905 a-d.

At least a portion of each vertical rail 905 a-d may be L-shaped in cross-section. The four vertical rails 905 a-d may be of same length. At least one of the vertical rails 905 a-d may have plurality of spaced apart apertures 915 along the length. The apertures 915 may be in the form of opening. These apertures 915 may facilitate mounting of components on the chassis 900.

The four horizontal rails 910A-D may also be of same length.

The internal volume of the chassis 900 may be cuboid. The chassis may be constructed of a construction material that is rigid. The chassis may be constructed of a 3D-printed titanium or other suitable alloy. The chassis 900 may be adapted to be used to support a satellite as described above.

Components such as but not limited to electromagnets can be built into the walls of chassis. Such arrangement does not compromise the interior volume of the satellite. In other words, by mounting the components of the satellite system into the walls, the interior space of the satellite is not occupied. A person skilled in the art may appreciate that such integration technique can provide significant advantage due to preservation of useful space/payload volume within the satellite.

Turning now to FIGS. 11A-11F, an electromagnet 1100 according to one preferred embodiment of the present invention will now be described. The electromagnet 1100 is a superconducting electromagnet (such as HTS superconducting electromagnet) suitable for use using the satellite 700 as described above with reference to FIG. 7 and also satellites 400, 500, 600 as described above with reference to FIGS. 4-6.

FIG. 11A shows a bottom cooling plate 1110 of the electromagnet 1100 according to one embodiment of the present invention. The bottom cooling plate 1110 and/or the top cooling plate 1100 may be of hexagonal shape and may be made out of a copper and may be 2 mm in thickness. As shown, the bottom cooling plate 1110 may comprise six 3 mm holes on 66 mm diameter.

A cylindrical magnet bore 1111, e.g. G10 cylindrical magnet bore may be provided as shown. The cylindrical magnet bore may have an outside diameter of about 10 mm and inside diameter of 8 mm.

Mild steel (magnetic) chamfered pole pieces 1112 a, 1112 b may be provided. A magnetic field sensor 1113 may be sandwiched between the two pole pieces 1112 a, 1112 b. Suitable support for the sensor and sensor leads may be provided.

Thermal link may be provided between top and bottom cooling plates. Current bus bar may be thermally anchored but may be electrically isolated by use of sapphire plates and may be attached by nylon screws.

FIG. 11B shows part of the electromagnet 1100 with a coil in the form of a double pancake (circular) coil 1120 in place. The coil may be made from approximately 100 m long, 3 mm wide, 50 μm thick Superpower wire or 2G YBCO HTS tape/wire producing approximately 60 mm diameter coils. Insulation sheet, e.g. G10 insulation sheet may be provided between the two coils and cooling plates. Coil windings may terminate on current bus. Thermal grease may be used in all thermal interfaces. The thermal grease may be Apeizon type N thermal grease. It should be noted that the coil need not be in a pancake or double pancake form. Many other suitable shapes of coil such as but not limited to rectangular, triangular etc. may be used. Similarly, the number of coils need not be two and can be one or greater than two.

FIG. 11C shows 1100 with top cooling plate 1130 in place. As shown in the cross-sectional views of the electromagnet 1100 (see FIG. 11D and FIG. 11E), yoke plates 1140 a 1140 b such as mild steel (magnetic) yoke plates may optionally be placed in the exterior surface at least one of the top cooling plate 1130 and the bottom cooling plate 1110. The yoke plates 1140 a, 1140 b may be screwed or lightly screwed to the cooling plates 1110, 1130.

As seen the cryocooling device in the form of a cryocooler 1150 may be mounted to the electromagnet 1100 for cooling. The cryocooler 1150 may be an internal Stirling or a pulse tube miniature tactical cryocooler. The inventors have found that use of such cryocoolers provides feasible cooling option in satellite environment. FIG. 12 shows one example of a cryocooler 1200 that may be used in the present invention.

Magnet assembly and cryocooler interface may be provided and intermediate cooling bus may be provided.

Securement means may be provided to ensure that the cooling plates are clamped securely for thermal contact with the coils. As shown in FIG. 11F brackets 1142 a, 1142 b, may be made from stainless steel may be used as securement means. The cryocooler is shown in FIG. 11F with reference numeral 1150.

In summary, the design of electromagnet 1100 according to the present invention may accommodate a 60 mm outer diameter double pancake coil 1120 a, 1120 b made from 3 mm HTS tape. The coil may use approximately 100 m of tape and may be dry-wound with no inter-turn insulation or may be embedded in a matrix with insulation. There may be an insulating sheet between the two pancake coils 1120 a, 1120 b, and on the outer face of each pancake to reduce electrical shorting. The coils 1120 a, 1120 b may be sandwiched between two copper plates that may be affixed to a cryocooler such as Ricor K508 N/K508 cryocooler. The coils 1120 a, 1120 b may be cooled by conduction through the copper plates.

The coils 1120 a, 1120 b may be wound on to an insulating mandrel or bore (preferably, G10 tube mandrel or G10 bore). A ferromagnetic core can optionally be placed within the bore to concentrate the magnetic field and increase the magnetic field density. Mild carbon steel or any other suitable material with a higher relative magnetic permeability and/or magnetic saturation point may be used, which has similar magnetic properties to iron.

Similarly, iron plates or mild steel plates 1140 a, 1140 b may be attached on the outside of the copper cooling plates 1110, 1130 to act as partial magnet yokes. The coil 1112 a, 1112 b would preferably be in two parts with a 1 mm gap so that a magnetic-field sensor 1113 can be placed in the gap. The addition of materials with high relative magnetic permeability and high magnetic saturation threshold may significantly enhance the magnetic field that can be achieved in the centre of the magnet for a given set of coils 1120 a 1120 b.

To further summarise, the electromagnet 1100 of the present invention may be suitable for use in a satellite 400, 500, 600, 700. The electromagnet may comprise: a top cooling plate 1130, a bottom cooling plate 1110, and coil such as a double pancake coil 1120 a, 1120 b sandwiched between the top cooling plate 1130 and the bottom cooling plate 1110. As the person skilled in the art may appreciate, the electromagnet 1100 may need a power source to function.

At least one of the top and bottom cooling plates 1130, 1110 may be hexagonal or substantially hexagonal is shape. At least one of the top and bottom cooling plates 1130. 1110 may be made from copper.

At least one of the top and bottom cooling plates 1130, 1110 may be 2 mm in thickness. The bottom cooling plate 1110 and/or top cooling plate 1130 may comprise six holes, preferably, six 3 mm holes on 66 mm diameter.

The electromagnet 1110 may comprise a cylindrical magnet bore. The cylindrical magnet bore may have an outside diameter of about 10 mm and inside diameter of 8 mm. The electromagnet 1110 may comprise core formed as two pole pieces 1112 a, 1112 b with a magnetic field sensor 1113 sandwiched between the pole pieces 1112 a, 1112 b.

Thermal link may be provided between top and bottom cooling plates 1130, 1110. In one embodiment, the coil 1120 a, 1120 b may be from approximately 100 m long, 3 mm wide, 50 μm thick Superpower wire or 2G YBCO HTS tape/wire. The coil 1120 a, 1120 b may have 60 mm outer diameter.

The coil 1120 a, 1120 b may use approximately 100 m of tape and may be dry-wound with no inter-turn insulation or embedded in a matrix with insulation. The insulation sheet (preferably G10 insulation sheet) may be provided between the coil (or double pancake coil) and cooling plates 1110, 1130 to reduce odds of electrical shorting.

The inner diameter of the double pancake coils may be approximately 10 mm. The thickness of the double pancake coils is approximately 3 mm. The outer diameter of the double pancake coil may be approximately 56 mm. The total number of turns of the wire or tape in the double pancake coil may be approximately 470. Total length of the wire or tape in the double pancake coil may be approximately 97.5 m. In one embodiment, the coil is wrapped around the cylindrical magnet bore.

For test setup, the windings of the coil 1120 a, 1120 b (coil windings) may terminate on a current bus.

The electromagnet 1110 may comprises at least one yoke plate 1140 a, 1140 b that may be attached to exterior surface of one or both of the top or bottom cooling plates 1130, 1110.

Thermal grease may be used or applied to all thermal interfaces. The thermal grease may be Apeizon type N thermal grease. The yoke plate(s) 1140 a, 1140 b may be a mild steel (magnetic) yoke plate(s).

The yoke plate(s) 1140 a, 1140 b may be screwed or lightly screwed to the top and bottom cooling plates 1130, 1110.

A cryocooler 1150, 1120 may be mounted to the electromagnet for cooling. The cryocooler 1150, 1200 may be an internal Stirling or a pulse tube miniature tactical cryocooler.

A securement means may be provided to clamp the top and bottom cooling plates 1130, 1110 securely for thermal contact with the coil (or double pancake coil) 1120 q, 1120 b. The securement means may be one or more brackets 1142 a, 1142 b may be made out of stainless steel.

The electromagnet 1100 may be adapted to be energised by electromagnetic flux injection device. The electromagnetic flux injection device may an electromagnetic flux pump. The electromagnetic flux pump may be a linear flux pump. The electromagnetic flux pump may be contactless. The electromagnet 1100 may be adapted to be used with the satellites as described above.

FIGS. 13A and B show an experimental set-up 1300. In FIG. 13A, electromagnet 1100 of the present invention can be seen at lower part of a lid 1310 of a vacuum chamber, which is for simulating space-like conditions on the Earth. The cryocooler 1150 can also be seen in FIG. 13A.

It will of course be realised that while the foregoing has been given by way of illustrative example(s) of the present invention, all such modifications and variations thereto as would be apparent to a person skilled in the art are deemed to fall within the broad scope and ambit of the various aspects if invention as is hereinbefore described and/or defined in the claims. 

1. A satellite comprising: a chassis; and a superconducting magnet control system mounted to or within at least a portion of said chassis for position control of said satellite, wherein, said superconducting magnet control system comprises: at least one superconducting electromagnet comprising at least one coil, and further comprising or mounted to at least one cooling element that is in thermal contact with said at least one coil; and at least one cryocooler; wherein, said at least one cryocooler is thermally coupled to said at least one cooling element to cool at least said at least one coil through said at least one cooling element by conduction cooling alone.
 2. The satellite as claimed in claim 1, wherein said at least one cryocooler is thermally coupled to said at least one cooling element for cooling said at least one cooling element so that when said at least one cooling element is at a lower temperature than said at least one coil, a transfer of heat through said at least one cooling element causes conduction cooling of at least said one coil.
 3. The satellite as claimed in claim 1, wherein said at least one cooling element is a cooling plate.
 4. The satellite as claimed in claim 3, wherein said cooling plate is a metallic or a non-metallic cooling plate.
 5. The satellite as claimed in claim 1, wherein said at least one superconducting electromagnet is a High Temperature Superconductor (HTS) electromagnet.
 6. The satellite as claimed in claim 1, wherein said at least one superconducting electromagnet comprises or is mounted to at least two cooling elements that are in thermal contact with said at least one coil, said cooling elements being a top cooling element and a bottom cooling element, wherein said at least one coil is sandwiched between said top cooling element and said bottom cooling element.
 7. The satellite as claimed in claim 1, wherein said at least one cryocooler is selected from a Stirling cryocooler, a pulse tube tactical cryocooler or a pulse tube miniature tactical cryocooler.
 8. The satellite as claimed in claim 1, wherein said superconducting magnet control system further comprises: at least one power source; and at least one control device; said at least one cryocooler being operatively coupled to said at least one control device and to said at least one power source, wherein based on a control signal from the control device said at least one cryocooler is configured to derive energy from said at least one power source to cool at least said at least one coil.
 9. The satellite as claimed in claim 8, wherein a set of reaction wheels that is operatively coupled to said at least one power source and said at least one control device.
 10. The satellite as claimed in claim 8, wherein said at least one control device derives energy from said at least one power source.
 11. The satellite as claimed in claim 8, wherein said at least one control device is configured to control at least one of a timing, a magnitude and a polarity of magnetic field in said at least one superconducting electromagnet.
 12. The satellite as claimed in claim 8, wherein said at least one control device is configured to control a timing of said at least one cryocooler to cool said at least one superconducting electromagnet.
 13. The satellite as claimed in claim 8, wherein said at least one power source comprises at least one solar panel.
 14. The satellite as claimed in claim 8, wherein said at least one power source is a battery or a capacitor.
 15. The satellite as claimed in claim 8, wherein said superconducting magnet control system further comprises at least one electromagnetic flux injection device that is operatively coupled to said at least one control device and is configured to derive energy from said at least one power source to energise said at least one superconducting electromagnet.
 16. The A satellite as claimed in claim 15, wherein said at least one electromagnetic flux injection device is controlled by said at least one control device to derive energy from said at least one power source to energise said at least one superconducting electromagnet.
 17. The satellite as claimed in claim 15, wherein said at least one electromagnetic flux injection device is an electromagnetic flux pump.
 18. The satellite as claimed in claim 17, wherein said at least one electromagnetic flux pump is a linear flux pump.
 19. The satellite as claimed in claim 17, wherein said at least one electromagnetic flux pump is contactless.
 20. The satellite as claimed in claim 19, wherein said at least one electromagnetic flux pump is a non-linear flux pump.
 21. The satellite as claimed in claim 17, wherein electromagnetic flux pump is configured to magnetise said at least one coil, said at least one coil being HTS coil or HTS tape.
 22. The satellite as claimed in claim 1, wherein said satellite has a mass of 500 kg or less.
 23. (canceled)
 24. A satellite comprising: a chassis; and a superconducting magnet control system mounted to or within at least a portion of the chassis for position control of said satellite, said superconducting magnet control system comprising: at least one superconducting electromagnet comprising at least one coil; at least one power source; at least one control device; at least one cryocooler; and at least one electromagnetic flux injection device; wherein, said at least one electromagnetic flux injection device is operatively coupled to said at least one control device and is configured to derive energy from said at least one power source to energise said at least one superconducting electromagnet, wherein, said at least one cryocooler being operatively coupled to said at least one control device and to said at least one power source, wherein based on a control signal from the control device said at least one cryocooler is configured to derive energy from said at least one power source to cool at least said at least one coil by conduction cooling alone, and wherein, said at least one superconducting electromagnet comprises or is mounted to at least one cooling element that is in thermal contact with said at least one coil and said at least one cryocooler is thermally coupled to said at least one cooling element to cool said at least one cooling element, thereby also cooling at least said at least one coil by conduction cooling alone.
 25. A superconducting electromagnet of a satellite, said superconducting electromagnet comprising at least one coil; wherein: said superconducting electromagnet further comprises or is mounted to a top cooling element and a bottom cooling element with said at least one coil being sandwiched between said top cooling element and said bottom cooling element, said at least one coil is in thermal contact with said top cooling element and said bottom cooling element, and wherein said top cooling element and said bottom cooling element are configured to be thermally coupled to at least one cryocooler to cool said at least one coil through said top and bottom cooling elements by conduction cooling alone.
 26. A superconducting magnet control system of a satellite for position control of said satellite, wherein, said superconducting magnet control system comprises: at least one superconducting electromagnet comprising at least one coil and further comprising or mounted to at least one cooling element that is in thermal contact with said at least one coil; and at least one cryocooler; wherein, said at least one cryocooler is thermally coupled to said at least one cooling element to cool at least said at least one coil through said at least one cooling element by conduction cooling alone.
 27. A method of cooling a superconducting electromagnet used in a satellite, the method comprising the steps of: providing a satellite comprising a superconducting magnet control system for position control of said satellite, said superconducting magnet control system having at least one superconducting electromagnet comprising at least one coil and further comprising or mounted to at least one cooling element that is in thermal contact with said at least one coil, thermally coupling at least one cryocooler to said at least one cooling element to cool at least said at least one coil through said at least one cooling element by conduction cooling alone.
 28. A spacecraft comprising: a chassis; and a superconducting magnet control system mounted to or within at least a portion of said chassis for position control of said spacecraft, wherein, said superconducting magnet control system comprises: at least one superconducting electromagnet comprising at least one coil and further comprising or mounted to at least one cooling element that is in thermal contact with said at least one coil; and at least one cryocooler; and wherein, said at least one cryocooler is thermally coupled to said at least one cooling element to cool at least said at least one coil through said at least one cooling element by conduction cooling alone.
 29. A spacecraft comprising: a chassis; and a superconducting magnet control system mounted to or within at least a portion of the chassis for position control of said spacecraft, said superconducting magnet control system comprising: at least one superconducting electromagnet comprising at least one coil; at least one power source; at least one control device; at least one cryocooler; and at least one electromagnetic flux injection device; wherein, said at least one electromagnetic flux injection device is operatively coupled to said at least one control device and is configured to derive energy from said at least one power source to energise said at least one superconducting electromagnet, wherein, said at least one cryocooler is operatively coupled to said at least one control device and to said at least one power source, wherein based on a control signal from the control device said at least one cryocooler is configured to derive energy from said at least one power source to cool at least said at least one coil by conduction cooling alone, and wherein, said at least one superconducting electromagnet comprises or is mounted to at least one cooling element that is in thermal contact with said at least one coil and said at least one cryocooler is thermally coupled to at least one cooling element to cool said at least one cooling element, thereby also cooling at least said at least one coil by conduction cooling alone.
 30. A superconducting electromagnet of a spacecraft, said superconducting electromagnet comprising at least one coil; wherein: said superconducting electromagnet further comprises or is mounted to a top cooling element and a bottom cooling element with said at least one coil being sandwiched between said top cooling element and said bottom cooling element, said at least one coil being in thermal contact with said top cooling element and said bottom cooling element, and wherein said top cooling element and said bottom cooling element are configured to be thermally coupled to at least one cryocooler to cool said at least one coil through said top and bottom cooling elements by conduction cooling alone.
 31. A superconducting magnet control system of a spacecraft for position control of said spacecraft, wherein, said superconducting magnet control system comprises: at least one superconducting electromagnet comprising at least one coil and further comprising or mounted to at least one cooling element that is in thermal contact with said at least one coil; and at least one cryocooler; wherein, said at least one cryocooler is thermally coupled to said at least one cooling element to cool at least said at least one coil through said at least one cooling element by conduction cooling alone. 